1. Field of the Invention
The present invention relates generally to earth-orbiting satellites, especially those in a geosynchronous orbit. A primary goal of the invention resides in increasing satellite payload capacity while moderating thermal transients in satellite equipment. Diurnal solar loading on the surfaces of a satellite presently has limited use east/west and aft radiators to control equipment which operates either discontinuously or at very high temperatures. This invention utilizes devices such as variable conduction head pipes (VC HP.sub.s), diode head pipes (DHP.sub.s) and phase change materials to sequentially connect and disconnect east/west/earth/anti-earth radiators to increase satellite heat dissipation capability. This invention permits higher thermal dissipation in current spacecraft design thus permitting the use of larger, more powerful, payloads without the need for costly redesign. Although the term geosynchronous satellite is used throughout the disclosure, it will be understood that the invention is applicable to any orbiting satellite with surfaces which are periodically exposed to direct solar illumination.
2. Description of the Prior Art
High power geosynchronous spacecraft are constrained in their power capability by, among other factors, the degree to which they can reject waste heat. Heat rejection can be increased by:
increasing the size of the satellite body, leading to heavier satellites, and thus higher costs; and PA1 use of deployable thermal radiators which intrinsically add cost and mass to the satellite. PA1 (1) connect the thermal load to a phase change management (PCM) material which is in turn connected to an intermittently available radiative surface with a variable conduction heat pipe (VCHP) or diode heat pipe (DHP); or PA1 (2) connect the thermal load to two or more opposed intermittently available radiative surfaces (east/west or earth/anti-earth) with sets of thermal conductors such as VCHPs or diode head pipes (DHPs) which will not conduct heat from a hot radiator to the internal equipment.
It was with knowledge of the foregoing state of the technology that the present invention has been conceived and is now reduced to practice. The actuation and deployment concept embodied by this invention is different from all of the devices reviewed above. Furthermore, this system can easily be implemented in existing as well as in new spacecraft without significant alterations in the design.